DOCSLIB.ORG

The Development of a Pulsed Plasma Thruster As a Solid Fuel Plasma Source for a High Power Helicon

Total Page：16

File Type：pdf, Size：1020Kb

Download full-text PDF Read full-text
The Development of a Pulsed Plasma Thruster As a Solid Fuel Plasma Source for a High Power Helicon The Development of a Pulsed Plasma Thruster as a Solid Fuel Plasma Source for a High Power Helicon Ian Kronheim Johnson A thesis submitted in partial fulfillment of the requirements for the degree of Master of Science The University of Washington 2011 Program authorized to offer degree: Aeronautics and Astronautics University of Washington Graduate School This is to certify that I have examined this copy of a master’s thesis by Ian Kronheim Johnson and have found that is it complete and satisfactory in all respects, and that any and all revisions required by the final examining committee have been made. Committee Members: Professor Robert Winglee, Department of Earth and Space Sciences, Chair Professor Tom Jarboe, Department of Aeronautics and Astronautics Date: The University of Washington ii In presenting this thesis in partial fulfillment of the requirements for a master’s degree at the University of Washington, I agree that the Library shall make its copies freely available for inspection. I further agree that extensive copying of this thesis is allowable only for scholarly purposes, consistent with “fair use” as prescribed in the U.S. Copyright Law. Any other reproduction for any purposes or by any means shall not be allowed without my written permission. Signature: Date: The University of Washington iii University of Washington Abstract The Development of a Pulsed Plasma Thruster as a Solid Fuel Plasma Source for a High Power Helicon Ian Kronheim Johnson Chair of the Supervisory Committee: Professor Robert Winglee Earth and Space Sciences As space exploration shifts to lower mass and lower cost missions, the need for improved on-board propulsion systems is growing. The High Power Helicon (HPH) experiment at the University of Washington is one new possibility in this field. The HPH is currently gas-fed, which creates a number of added concerns during operation. For this reason, a Pulsed Plasma Thruster (PPT) was developed to act as a solid fuel source for the Helicon thruster. The PPT created was 50% smaller in size than the EO-1 satellite PPT [Zakrzwski] and produced electron densities on the order of 1018 m-3, 20 cm downstream of the Teflon surface at 43 J. The thruster had an estimated impulse bit of 138 µN-sec and a specific impulse of 1413 seconds while ablating 44 µg of Teflon per pulse at 4.4% efficiency. The HPH system currently uses a helicon antenna to ionize and accelerate the gas particles with high efficiency. Therefore, the objective of the PPT for the HPH operating system was to produce large numbers of neutral particles traveling at low to medium velocities. This is in comparison to most PPT investigations where higher efficiencies and thrusts are desired. The PPT was successfully fired under vacuum while operated with the HPH magnetic field and is ready to be fully integrated and tested with the HPH system. The University of Washington iv Table of Contents 1 Introduction..............................................................................................................................................1 1.1 Current Propulsion Concepts ...................................................................................................1 1.1.1 Electrostatic Propulsion.....................................................................................................2 1.1.2 Electromagnetic Propulsion.............................................................................................3 1.1.3 Electrothermal Propulsion................................................................................................5 1.1.4 Chemical and Electrical Comparison............................................................................7 1.2 High Power Helicon Thruster...................................................................................................8 1.3 Helicon Thruster Design..........................................................................................................10 1.3.1 Gas vs. Solid Fed Helicon Thruster.............................................................................11 1.4 Pulsed Plasma Thruster...........................................................................................................12 1.4.1 PPT Operation .....................................................................................................................13 1.4.2 PPT Plume .............................................................................................................................15 1.4.3 PPT Flight History..............................................................................................................16 1.5 Outline and Research Goals.................................................................................................... 19 2 Experimental Setup and Electronics Design............................................................................21 2.1 Vacuum System ...........................................................................................................................21 2.2 Control and Data Acquisition ................................................................................................22 2.3 Electronic Control Systems .................................................................................................... 23 2.3.1 Starter Power Processing Unit .....................................................................................23 2.3.2 Main Discharge Power Processing Unit ................................................................... 24 2.4 HPH Base Magnets .....................................................................................................................27 3 Diagnostics.............................................................................................................................................28 3.1 Stangenes Probe..........................................................................................................................28 3.2 Voltage Dividers and the High Voltage Probe ................................................................28 3.3 Langmuir Probes.........................................................................................................................29 3.3.1 Standard Double Probe Theory ...................................................................................30 3.3.2 Probe Construction and Description ......................................................................... 34 3.4 CCD Camera................................................................................................................................... 35 4 Theoretical Models and Data Reduction ...................................................................................36 5 PPT Iterations and Discussion....................................................................................................... 38 5.1 PPT-1................................................................................................................................................38 5.1.1 Spark Plug – PPTv1.0........................................................................................................ 38 The University of Washington v 5.1.2 Spark Wire – PPTv1.1 & 1.2...........................................................................................39 5.2 PPT - 2..............................................................................................................................................41 5.2.1 No Insulation - PPTv2.0................................................................................................... 41 5.2.2 Insulation - PPTv2.1..........................................................................................................42 5.2.3 Electron Density On-axis................................................................................................. 43 5.2.4 Plume Width.........................................................................................................................47 5.2.5 Predicted Electron Densities at 10 cm height........................................................54 5.2.6 Time-of-Flight Velocity Data .........................................................................................54 5.2.7 Mass Ablation Estimates................................................................................................. 57 5.2.8 PPT Performance Estimates..........................................................................................58 5.2.9 Effect of an Applied Magnetic Field............................................................................59 6 Summary.................................................................................................................................................63 7 Bibliography..........................................................................................................................................65 The University of Washington vi List of Figures Figure 1.1. An electrostatic gridded ion thruster schematic (left) [XIPS] and a xenon ion discharge from the NSTAR ion thruster (right) [Ion Propulsion]. .................................2 Figure 1.2. The MPD schematic (left) [Jahn and Choueiri] and (right) a NASA MPD thruster experiment in operation (right) [MPD: NASA Facts]. ...............................................4 Figure 1.3. The Hall thruster schematic (left) [Jahn and Choueiri] and a NASA Hall thruster firing with xenon propellant (right) [Electric Propulsion]. ....................................5
Load more

Recommended publications
  • DESIGN and DEVELOPMENT of a 30-Ghz MICROWAVE ELECTROTHERMAL THRUSTER
    The Pennsylvania State University The Graduate School College of Engineering DESIGN AND DEVELOPMENT OF A 30-GHz MICROWAVE ELECTROTHERMAL THRUSTER A Thesis in Aerospace Engineering by Erica E. Capalungan ©2011 Erica E. Capalungan Submitted in Partial Fulfillment of the Requirements for the Degree of Master of Science August 2011 The thesis of Erica E. Capalungan was reviewed and approved* by the following: Michael M. Micci Professor of Aerospace Engineering Director of Graduate Studies Thesis Advisor Sven G. Bilén Associate Professor of Engineering Design, Electrical Engineering, and Aerospace Engineering George A. Lesieutre Professor of Aerospace Engineering Head of the Department of Aerospace Engineering *Signatures are on file in the Graduate School. ii ABSTRACT Research has been conducted on the microwave electrothermal thruster at The Pennsylvania State University since the 1980’s. Each subsequent thruster incorporated modifications that resulted in improvements in thruster performance compared to previous generations. Operational frequencies evaluated thus far include 2.45 GHz, 7.5 GHz, 8 GHz, and 14.5 GHz. As each thruster increased in operational frequency, plasmas have been ignited with successively lower amounts of input power. With higher frequency and lower power requirements, the physical sizes of the thruster and the power supply have been reduced. Decreased size results in a lighter propulsion system, which is ideal for space missions. This thesis concerns the design and development of a thruster operating at 30 GHz. Electromagnetic modeling was used in the design of the thruster to determine the optimal input antenna size and length. A 2.4-mm antenna size was chosen with a length that is flush with the bottom of the cavity.
    [Show full text]
  • 1 Iac-06-C4.4.7 the Innovative Dual-Stage 4-Grid Ion
    IAC-06-C4.4.7 THE INNOVATIVE DUAL-STAGE 4-GRID ION THRUSTER CONCEPT – THEORY AND EXPERIMENTAL RESULTS Cristina Bramanti, Roger Walker, ESA-ESTEC, Keplerlaan 1, 2201 AZ Noordwijk, The Netherlands [email protected], Roger. Walker @esa.int Orson Sutherland, Rod Boswell, Christine Charles Plasma Research Laboratory, Research School of Physical Sciences and Engineering, The Australian National University, Canberra, ACT 0200, Australia [email protected], [email protected], [email protected]. David Fearn EP Solutions, 23 Bowenhurst Road, Church Crookham, Fleet, Hants, GU52 6HS, United Kingdom [email protected] Jose Gonzalez Del Amo, Marika Orlandi ESA-ESTEC, Keplerlaan 1, 2201 AZ Noordwijk, The Netherlands [email protected], [email protected] ABSTRACT A new concept for an advanced “Dual-Stage 4-Grid” (DS4G) ion thruster has been proposed which draws inspiration from Controlled Thermonuclear Reactor (CTR) experiments. The DS4G concept is able to operate at very high specific impulse, power and thrust density values well in excess of conventional 3-grid ion thrusters at the expense of a higher power-to-thrust ratio. A small low-power experimental laboratory model was designed and built under a preliminary research, development and test programme, and its performance was measured during an extensive test campaign, which proved the practical feasibility of the overall concept and demonstrated the performance predicted by analytical and simulation models. In the present paper, the basic concept of the DS4G ion thruster is presented, along with the design, operating parameters and measured performance obtained from the first and second phases of the experimental campaign.
    [Show full text]
  • Performance Optimization Criteria for Pulsed Inductive Plasma Acceleration Kurt A
    IEEE TRANSACTIONS ON PLASMA SCIENCE, VOL. 34, NO. 3, JUNE 2006 945 Performance Optimization Criteria for Pulsed Inductive Plasma Acceleration Kurt A. Polzin and Edgar Y. Choueiri Abstract—A model of pulsed inductive plasma thrusters con- sisting of a set of coupled circuit equations and a one-dimensional momentum equation has been nondimensionalized leading to the identification of several scaling parameters. Contour plots representing thruster performance (exhaust velocity and effi- ciency) were generated numerically as a function of the scaling parameters. The analysis revealed the benefits of underdamped current waveforms and led to an efficiency maximization criterion that requires the circuit’s natural period to be matched to the acceleration timescale. It is also shown that the performance increases as a greater fraction of the propellant is loaded nearer to the inductive acceleration coil. Index Terms—Acceleration modeling, nondimensional scaling parameters, pulsed inductive acceleration, pulsed plasma acceler- ation. I. INTRODUCTION ULSED inductive plasma accelerators are spacecraft P propulsion devices in which energy is stored in a capacitor and then discharged through an inductive coil. The device is Fig. 1. (a) Conceptual schematic (from [5]) and (b) general lumped element circuit model (after [5]) of a pulsed inductive accelerator. electrodeless, inducing a current in a plasma located near the face of the coil. The propellant is accelerated and expelled at a high exhaust velocity ( (10 km/s)) by the Lorentz force arising from the interaction of the plasma current and the In PPT research, there has been substantial effort devoted induced magnetic field [see Fig. 1(a) for a thruster schematic].
    [Show full text]
  • Latin Derivatives Dictionary
    Dedication: 3/15/05 I dedicate this collection to my friends Orville and Evelyn Brynelson and my parents George and Marion Greenwald. I especially thank James Steckel, Barbara Zbikowski, Gustavo Betancourt, and Joshua Ellis, colleagues and computer experts extraordinaire, for their invaluable assistance. Kathy Hart, MUHS librarian, was most helpful in suggesting sources. I further thank Gaylan DuBose, Ed Long, Hugh Himwich, Susan Schearer, Gardy Warren, and Kaye Warren for their encouragement and advice. My former students and now Classics professors Daniel Curley and Anthony Hollingsworth also deserve mention for their advice, assistance, and friendship. My student Michael Kocorowski encouraged and provoked me into beginning this dictionary. Certamen players Michael Fleisch, James Ruel, Jeff Tudor, and Ryan Thom were inspirations. Sue Smith provided advice. James Radtke, James Beaudoin, Richard Hallberg, Sylvester Kreilein, and James Wilkinson assisted with words from modern foreign languages. Without the advice of these and many others this dictionary could not have been compiled. Lastly I thank all my colleagues and students at Marquette University High School who have made my teaching career a joy. Basic sources: American College Dictionary (ACD) American Heritage Dictionary of the English Language (AHD) Oxford Dictionary of English Etymology (ODEE) Oxford English Dictionary (OCD) Webster’s International Dictionary (eds. 2, 3) (W2, W3) Liddell and Scott (LS) Lewis and Short (LS) Oxford Latin Dictionary (OLD) Schaffer: Greek Derivative Dictionary, Latin Derivative Dictionary In addition many other sources were consulted; numerous etymology texts and readers were helpful. Zeno’s Word Frequency guide assisted in determining the relative importance of words. However, all judgments (and errors) are finally mine.
    [Show full text]
  • Electrostatic Ion Thrusters for Space Debris Removal
    Theoretical Physics Electrostatic ion thrusters for space debris removal Oscar Larsson, [email protected] Gustav Hedengren, [email protected] SA114X Degree Project in Engineering Physics, First Level Department of Theoretical Physics Royal Institute of Technology (KTH) Supervisor: Christer Fuglesang October 28, 2018 Abstract The current levels of space debris are critical and actions are needed to prevent collisions. In this paper it is examined whether an electrostatic ion thruster can be powerful enough to slow down the debris in a sufficient manner. Furthermore, it is looked into whether the process can be repeated for a significant number of pieces by maneuvering between them. We conclude that the removal process seems possible although some improvements are needed. Maneuvering is costly but despite conservative assumptions, we estimate that about 800 pieces can be removed in one journey made by a satellite weighing ten tonnes of which nine are xenon. Abstract Den nuvarande nivån av rymdskrot är kritisk och åtgärder krävs för att förhindra kollisioner. I denna artikel undersöks det huruvida en elektrostatisk jonkanon är kraftfull nog för att bromsa skrot tillräckligt. Fortsättningsvis undersöks det om denna process är effectiv nog för att återupprepas för ett betydande antal bitar, inklusive manövrering bitarna emellan. Vi drar slutsatsen att processen verkar möjlig att genomföra även om vissa förbättringar behövs. Manövreringen är kostsam men trots konservativa antaganden uppskattar vi att ungefär 800 bitar kan tas ned under en resa av en satellit med vikt 10 ton varav nio ton är xenon. Contents 1 Introduction 2 2 Background 2 2.1 The space debris problem . .2 2.2 The Dual-stage 4 grid ion thruster (DS4G) .
    [Show full text]
  • Scaling Laws for Electromagnetic Pulsed Plasma Thrusters
    INSTITUTE OF PHYSICS PUBLISHING PLASMA SOURCES SCIENCE AND TECHNOLOGY Plasma Sources Sci. Technol. 10 (2001) 395–405 PII: S0963-0252(01)24935-4 Scaling laws for electromagnetic pulsed plasma thrusters J K Ziemer1 and E Y Choueiri Electric Propulsion and Plasma Dynamics Laboratory (EPPDyL), Applied Physics Group, Mechanical and Aerospace Engineering Department, Princeton University, Princeton, NJ 08544, USA Received 23 March 2001, in final form 7 May 2001 Published 5 June 2001 Online at stacks.iop.org/PSST/10/395 Abstract The scaling laws of pulsed plasma thrusters operating in the predominantly electromagnetic acceleration mode (EM-PPT) are investigated theoretically and experimentally using gas-fed pulsed plasma thrusters. A fundamental characteristic velocity that depends on the inductance per unit length and the square root of the capacitance to the initial inductance ratio is identified. An analytical model of the discharge current predicts scaling laws in which the propulsive efficiency is proportional to the EM-PPT performance scaling number, defined here as the ratio of the exhaust velocity to the EM-PPT characteristic velocity. The importance of the effective plasma resistance in improving the propulsive performance is shown. To test the validity of the predicted scaling relations, the performance of two gas-fed pulsed plasma thruster designs (one with coaxial electrodes and the other with parallel-plate electrodes), was measured under 70 different operating conditions using an argon plasma. The measurements demonstrate that the impulse bit scales linearly with the integral of the square of the discharge current as expected for an electromagnetic accelerator. The measured performance scaling is shown to be in good agreement with the theoretically predicted scaling.
    [Show full text]
  • Electric Propulsion 4
    Technology Forum on Small Body Scientific Exploration – 4th Meeting of the NASA Small Bodies Assessment Group Michael Patterson NASA Glenn Research Center John Brophy Jet Propulsion Laboratory California Institute of Technology January 24, 2011 National Aeronautics and Space Administration www.nasa.gov 1. In-Space Propulsion Overview 2. Current In-Space Investments for Robotic Missions 3. Description of Electric Propulsion 4. Current Technology Investments and Development Status 2 3 Flight System Development Future Applications Non-toxic propellant-based propulsion systems (LOX, LH2, LCH4, Ethanol) Main Propulsion Systems (MPS) Propulsion systems for orbit transfer, orbit injection, spacecraft maneuvering, landing, and ascent Development of component technologies (igniters, exciters, injectors , combustion chambers, nozzles) for non-toxic propellants Current Focus Areas • LOX/LCH4 propulsion technologies for main propulsion systems (MPS), reaction control systems (RCS), and propellant storage and distribution • Development of Orion Service Module propulsion system Reaction Control Systems (RCS) • Preliminary trade studies, requirements development, and planning for Altair ascent propulsion subsystem Research and Technology Development Materials Evaluation Combustion Diagnostics Ignition Performance Propellant Properties 4 Future Applications • Next generation ion propulsion system for deep-space science missions (NEXT) • Propulsion systems for Earth orbital applications including satellite servicing, repositioning, re-boost, orbital
    [Show full text]
  • Evolution of the AIRBUS DS Gmbh Radio Frequency Ion Thruster Family
    Evolution of the AIRBUS DS GmbH Radio Frequency Ion Thruster Family IEPC-2015- 90830 Presented at Joint Conference of 30th International Symposium on Space Technology and Science 34th International Electric Propulsion Conference and 6th Nano-satellite Symposium, Hyogo-Kobe, Japan July 4 – 10, 2015 H. J. Leiter1, Ch.Altmann2, R.Kukies3, J.Kuhmann4, J.-P. Porst5 Airbus DS GmbH, Lampoldshausen, Germany Marcel Berger6 and Michael Rath7 Airbus DS GmbH, Bremen, Germany Abstract: The Airbus DS Radio Frequency Ion Thruster "RIT" family and its evolution is presented. The first member of the family was RIT-10. Over the years the engine was carefully evolved. The recent RIT-10 EVO represents the state-of-the art in the 500W class. Taking benefit of the excellent scalability of the RIT technology the family grew in both directions and a thrust range from view micro Newtons to more than 200mN is covered. RIT-μX is the smallest engine for the low Milli- and Micro-Newton regime. In the 5kW class the RIT 2X series replaces RIT-22. The basics of the radio-frequency ion thruster technology are provided. It is a flight proven technology: RIT-10 marked two milestones in the history of electric propulsion. Following a brief historical recall of the RIT-family the three thrusters RIT-μX, RIT-10 EVO and the RIT 2X series are presented together with their embedding systems. I. Introduction lectric propulsion (EP) has become state of the art for commercial geostationary satellites. More and more Eoperators realize the advantages of EP for North South Station keeping and thanks to sufficient electric power on telecom platforms also electric orbit transfer will soon become standard.
    [Show full text]
  • Future Directions for Electric Propulsion Research
    aerospace Article Future Directions for Electric Propulsion Research Ethan Dale * , Benjamin Jorns and Alec Gallimore Department of Aerospace Engineering, University of Michigan, Ann Arbor, MI 48105, USA; [email protected] (B.J.); [email protected] (A.G.) * Correspondence: [email protected] Received: 7 July 2020; Accepted: 17 August 2020; Published: 20 August 2020 Abstract: The research challenges for electric propulsion technologies are examined in the context of s-curve development cycles. It is shown that the need for research is driven both by the application as well as relative maturity of the technology. For flight qualified systems such as moderately-powered Hall thrusters and gridded ion thrusters, there are open questions related to testing fidelity and predictive modeling. For less developed technologies like large-scale electrospray arrays and pulsed inductive thrusters, the challenges include scalability and realizing theoretical performance. Strategies are discussed to address the challenges of both mature and developed technologies. With the aid of targeted numerical and experimental facility effects studies, the application of data-driven analyses, and the development of advanced power systems, many of these hurdles can be overcome in the near future. Keywords: electric propulsion; Hall effect thruster; gridded ion thruster; electrospray; magnetic nozzle; pulsed inductive thruster 1. Introduction The use of electric propulsion (EP) for space applications is currently undergoing a rapid expansion. There are hundreds of operational spacecraft employing EP technologies with industry projections showing that nearly half of all commercial launches in the next decade will have a form of electric propulsion. In light of their widespread use, the thruster types that have fueled this expansion—moderately-powered (1–20 kW) Hall effect, electrothermal, and ion thrusters—arguably have now achieved “mature” operational status.
    [Show full text]
  • Liquid Pulsed Plasma Thruster Plasma Plume Investigation and MR-SAT Cold Gas Propulsion System Performance Analysis
    Scholars' Mine Masters Theses Student Theses and Dissertations Summer 2017 Liquid pulsed plasma thruster plasma plume investigation and MR-SAT cold gas propulsion system performance analysis Jeremiah Daniel Hanna Follow this and additional works at: https://scholarsmine.mst.edu/masters_theses Part of the Aerospace Engineering Commons Department: Recommended Citation Hanna, Jeremiah Daniel, "Liquid pulsed plasma thruster plasma plume investigation and MR-SAT cold gas propulsion system performance analysis" (2017). Masters Theses. 7945. https://scholarsmine.mst.edu/masters_theses/7945 This thesis is brought to you by Scholars' Mine, a service of the Missouri S&T Library and Learning Resources. This work is protected by U. S. Copyright Law. Unauthorized use including reproduction for redistribution requires the permission of the copyright holder. For more information, please contact [email protected]. i LIQUID PULSED PLASMA THRUSTER PLASMA PLUME INVESTIGATION AND MR-SAT COLD GAS PROPULSION SYSTEM PERFORMANCE ANALYSIS by JEREMIAH DANIEL HANNA A THESIS Presented to the Faculty of the Graduate School of the MISSOURI UNIVERSITY OF SCIENCE AND TECHNOLOGY In Partial Fulfillment of the Requirements for the Degree MASTER OF SCIENCE IN AEROSPACE ENGINEERING 2017 Approved by Dr Joshua Rovey, Advisor Dr Henry Pernicka, Co-Advisor Dr David Riggins ii iii ABSTRACT The plasma plume produced by a liquid pulsed plasma thruster is investigated using a Langmuir triple probe and a nude Faraday probe. The Langmuir triple probe failed to produce results which are suspected caused by the presence of ionic liquid in the plume resulting in shorting of the probe. The nude Faraday probe is able to record the ion current density which revealed a high level of inconsistency in the plasma plume.
    [Show full text]
  • Parametric Studies of the Hall Thruster at Soreq
    IAEC - Parametric Studies of the Hall Thntster at Soreq An elec IL9906633 initiated at Soreq Parametric a few years ago, aiming at the research and development of Studies of advanced Hall thrusters for various space applications. The Hall thruster accelerates a plasma jet by an axial electric the Hall field and an applied radial magnetic field in an annular Thruster at ceramic channel. A relatively large current density (> 0.1 A/cm") can be obtained, since the acceleration mechanism Soreq is not limited by space charge effects. Such a device can be used as a small rocket engine onboard spacecraft with the J. Ashkenazy, advantage of a large jet velocity compared with Y. Raitses and conventional rocket engines (10,000-30,000 m/s v.s G. Appelbaum 2,000-4,800 m/s). An experimental Hall thruster was constructed at Soreq and operated under a broad range of operating conditions and under various configurational /. Summary variations. Electrical, magnetic and plasma diagnostics, as well as accurate thrust and gas flow rate measurements, have been used to investigate the dependence of thruster behavior on the applied voltage, gas flow rate, magnetic field, channel geometry and wall material. Representative results highlighting the major findings of the studies conducted so far are presented. 2. Introduction The Hall current plasma thruster is a crossed field discharge device in which quasi-neutral plasma is accelerated by axial electric and radial magnetic fields in an annular channel. Thrust is generated as a reaction to the momentum carried by the plasma jet emerging from the channel's open end.
    [Show full text]
  • History and Current Status of the Microwave Electrothermal Thruster
    Progress in Propulsion Physics 1 (2009) 425-438 DOI: 10.1051/eucass/200901425 © Owned by the authors, published by EDP Sciences, 2009 HISTORY AND CURRENT STATUS OF THE MICROWAVE ELECTROTHERMAL THRUSTER M. M. Micci, S. G. Bil‚en, and D. E. Clemens The microwave electrothermal thruster (MET) heats a propellant by means of a free-§oating microwave-generated plasma within a microwave resonant cavity followed by a gasdynamic nozzle expansion. The MET design is detailed along with computational electromagnetic modeling of various resonant cavities. Performance of thrusters operating at mi- crowave frequencies of 2.45, 7.5, and 14 GHz at power levels between 10 and 2500 W is discussed. Measurements of thruster electromagnetic interference (EMI) indicate that the thruster emits a very low level of EMI compared to other electric propulsion devices. Laboratory meas- urements of exhaust velocity, thrust, and speci¦c impulse for several propellants are presented. 1 INTRODUCTION The MET is an electric propulsion device that uses microwaves to heat a gaseous propellant. Microwave energy is fed into an electromagnetic resonant cavity to ignite and sustain a free-§oating plasma, which heats the propellant gas. The heated propellant is then accelerated through a gasdynamic nozzle and exhausted to generate thrust. The heating mechanism is similar to that of an arcjet, which utilizes an arc discharge formed between two electrodes to heat a propellant gas. The main di¨erence is that the MET plasma is free-§oating and thus the system does not su¨er from the lifetime-limiting electrode erosion problems that are characteristic of the arcjet.
    [Show full text]